This disclosure generally relates to hybrid electrical power sources having two or more electrical energy sources that supply energy to a connected load. In particular, the technology disclosed herein relates to hybrid electrical power sources comprising one or more batteries and one or more electric generators driven by internal combustion engines or gas turbine engines.
Some aircraft have electrically powered propulsion systems (hereinafter “electric aircraft”). In such aircraft, electric motors convert electrical power into mechanical power for use by the propulsion system. For example, an electric motor may turn one or more propellers on the aircraft to provide thrust. An electric aircraft may take various forms. For example, the electric aircraft may be an aircraft, a rotorcraft, a helicopter, a quadcopter, an unmanned aerial vehicle, or some other suitable type of aircraft.
When electric motors are used for propulsion of the aircraft, electrical energy may be supplied by a power source. For instance, electrical energy may be supplied using a battery system. The load on the battery system or other power source is an important consideration for the design and manufacturing of the aircraft. For example, the amount of electrical energy used by the electric motor to move the aircraft during various stages of flight may be important. Electric motors that use battery systems may require the battery to be recharged after a specified amount of time, distance, electrical energy use, or a combination thereof.
Some electric aircraft have a hybrid electric power architecture (hereinafter “hybrid electric aircraft”) in which at least two different types of power sources are connected in parallel to a load. The electrical energy sources will often have different electrical characteristics. For example, the electrical energy sources may be a battery and an electric generator driven by an internal combustion engine or a gas turbine engine. The battery supplies electrical power to an electric motor that is arranged to convert electrical power into mechanical power for use by the propulsion system of the aircraft.
In the case of a battery-equipped hybrid electric aircraft, the battery voltage cannot be actively controlled. Battery voltage is determined by the number of cells, type of cells, battery state of charge (SOC), loading and other factors. It is necessary to control power flow to and from the battery. Rate of charge or discharge of the battery is important and should be controlled to avoid thermal runaway.
For hybrid electric aircraft, the batteries are large and designed to provide a large amount of power for the purpose of propulsion. The batteries are paralleled with other power sources, such as electric generators. In one implementation, the battery is connected to a high-voltage direct-current (HVDC) bus, which is also supplied by the generator source(s). As used in the aerospace industry and herein, the term “high voltage” in the context of direct current means any DC voltage higher than 500 VDC. Such DC high voltage is typically derived from rectification of three-phase 230 VAC power.
There are no existing solutions for active power flow control and battery power management control for hybrid electric aircraft that employ batteries with hundreds of kilowatt-hours of energy. A system and method for tightly controlling the power flow to and from the battery at the HVDC connection is wanted.